1. Field of the Invention
The present invention relates generally to a stator vane in an industrial gas turbine engine, and more specifically to a large chord turbine vane with a cooling circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, a turbine section includes a plurality of stages of stator vanes and rotor blades to extract mechanical energy from the hot gas flow passing through the turbine. The efficiency of the turbine, and therefore of the engine, can be increased by increasing the turbine inlet temperature of the gas flow from the combustor. However, the temperature is limited to the material properties of the first stage turbine airfoils—the stator vanes and rotor blades—since the first stage airfoils are exposed to the hottest gas flow.
Passing cooling air through the airfoils can also allow for a higher gas flow temperature since the cooled airfoils can be exposed to higher temperatures. Complex convection and film cooling circuits have been proposed in the prior art to maximize the cooling effectiveness of the internal cooling circuits. Increasing the cooling ability while using less cooling air will provide higher efficiency.
In the prior art (U.S. Pat. No. 5,488,825 issued to Davis et al on Feb. 6, 1996 and entitled GAS TURBINE VANE WITH ENHANCED COOLING) large turbine vanes with cooling circuits, the vane airfoil is coupled with a large I.D. and O.D. endwall structure. Frequently an uneven gas temperature profile will enter the turbine stage and the airfoil endwalls are exposed to different gas temperature loading conditions. Also, due to an uneven cooling for the airfoil versus the endwalls, it is very difficult for a cooling circuit design to achieve a long balanced LCF life for a large turbine vane. The Davis et al patent show one prior art vane cooling circuit for a third stage vane of an industrial gas turbine engine which includes a 5-pass aft flowing serpentine cooling circuit with built-in cooling flow modulation device that can be used in a large chord turbine vane cooling circuit.
However, the stator vane cooling circuit of the Davis et al patent has some disadvantages. For the vane trailing edge OD fillet region, due to inadequate cooling for the junction of the airfoil trailing edge fillet versus the endwall location, the vane aft fillet region experiences a low LCF (low cycle fatigue) life. Also, at the vane trailing edge fillet location, a higher heat transfer coefficient or heat load onto the downstream fillet location exists due to the trailing edge wake effect. On top of a higher heat load onto the airfoil fillet location due to the stress concentration issue, the cooling hole for the airfoil trailing edge OD section cannot be located high enough into the vane OD section fillet region to provide proper convective cooling. Cooling of this particular airfoil trailing edge fillet region becomes especially difficult. As the turbine inlet temperature increases with improvements in the technologies of gas turbine engines, larger rotor blades and stator vanes are needed to extract the higher available energy from the hot gas flow. Thus, there is a need in the art of industrial gas turbine engines for larger turbine vanes with increased cooling abilities in order to withstand the larger blade size and higher heat loads.
It is therefore an object of the present invention to provide for a large stator turbine vane with an improved cooling circuit.
It is another object of the present invention to provide for a turbine stator vane with a 5 pass serpentine flow cooling circuit with built-in flow modulation that can be used in a large chord turbine vane cooling circuit.